U.S. patent number 5,081,832 [Application Number 07/724,476] was granted by the patent office on 1992-01-21 for high efficiency, twin spool, radial-high pressure, gas turbine engine.
This patent grant is currently assigned to Rolf Jan Mowill. Invention is credited to Rolf J. Mowill.
United States Patent |
5,081,832 |
Mowill |
January 21, 1992 |
High efficiency, twin spool, radial-high pressure, gas turbine
engine
Abstract
A low cost, low fuel consumption, twin-spool turbine engine with
external power take-off essentially exclusively from the high
pressure (H.P.) spool uses a radial in-flow turbine and a
centrifugal compressor in the H.P. spool and a flow-optimized L.P.
spool to achieve pressure ratios above 12/1 for simple cycle
engines. Intercooled and recuperated versions are included. The
turbine engine according to this invention is ideally suited for
applications up to about 1000 KW and will provide some 20% less
fuel consumption in this power range compared with conventional
designs.
Inventors: |
Mowill; Rolf J. (0386 Oslo, 3,
NO) |
Assignee: |
Mowill; Rolf Jan (Oslo,
NO)
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Family
ID: |
27049235 |
Appl.
No.: |
07/724,476 |
Filed: |
June 28, 1991 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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488138 |
Mar 5, 1990 |
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Current U.S.
Class: |
60/792 |
Current CPC
Class: |
F02C
3/36 (20130101); F02C 3/103 (20130101); F02C
3/09 (20130101); F02B 1/04 (20130101); F02B
2075/027 (20130101) |
Current International
Class: |
F02C
3/00 (20060101); F02C 3/09 (20060101); F02C
3/36 (20060101); F02C 3/10 (20060101); F02B
75/02 (20060101); F02B 1/04 (20060101); F02B
1/00 (20060101); F02C 003/05 () |
Field of
Search: |
;60/39.161,39.17,39.36
;415/179 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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0247984 |
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Dec 1987 |
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EP |
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748208 |
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Oct 1944 |
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DE2 |
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833739 |
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Mar 1952 |
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DE |
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2812237 |
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Oct 1978 |
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DE |
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1246455 |
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Oct 1960 |
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FR |
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283198 |
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May 1952 |
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CH |
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668834 |
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Mar 1952 |
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GB |
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720436 |
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Dec 1954 |
|
GB |
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Other References
Shepherd, D. G. Principles of Turbomachinery. New York: The
Macmillan Co., 1956, pp. 282-289. .
Csanady, G. T. Turbomachines New York: McGraw-Hill Book Co., 1964,
pp. 14-23. .
Cox, H. R., "Gas Turbine Principles and Practice," Van Nostrand,
pp. 2-26 1955. .
Traegor, E. E., "Aircraft Gas Turbine Technology", p. 10,
1970..
|
Primary Examiner: Bertsch; Richard A.
Assistant Examiner: Thorpe; Timothy S.
Attorney, Agent or Firm: Finnegan, Henderson, Farabow,
Garret & Dunner
Parent Case Text
This application is a continuation of application Ser. No.
07/488,138, filed Mar. 5, 1990 now abandoned.
Claims
What is claimed is:
1. A high-efficiency, twin spool gas turbine engine for driving an
external load, the engine comprising:
a high pressure spool having shaft means, and a centrifugal
compressor and a radial in-flow turbine mounted on said shaft
means, said radial in-flow turbine having an expansion ratio of
about 4/1 to 8/1;
means connected to said shaft means for driving essentially the
entire external load from said high pressure spool shaft means;
combustion means associated with said high pressure spool for
receiving compressed air from said centrifugal compressor, for
combusting fuel using the compressed air, and for delivering the
combustion gases and all the combustion heat energy to said radial
in-flow turbine under all engine load conditions;
a low pressure spool having shaft means; and low pressure
compressor means and low pressure turbine means mounted on said low
pressure spool shaft means, said low pressure spool being rotatably
independent from said high pressure spool and being essentially
load-free;
first means for flow interconnecting said low pressure compressor
means and said high pressure centrifugal compressor for supplying
precompressed air thereto; and
second means for flow interconnecting said high pressure radial
in-flow turbine means and said low pressure turbine means for
supplying partially expanded combustion gases thereto, wherein the
expansion ratio across the radial in-flow turbine is greater than
the expansion ratio across the low pressure turbine.
2. The twin spool gas turbine engine as in claim 1 further
including intercooler means associated with said first flow
connection means for cooling said pre-compressed air.
3. The twin spool gas turbine engine as in claim 2 wherein the
pressure ratio of said low pressure compressor means is less than
or equal to the pressure ratio of said high pressure centrifugal
compressor.
4. The twin spool gas turbine engine as in claim 1 further
including recuperator means for extracting heat values from said
combustion gases exhausted from said low pressure turbine means and
preheating the compressed air delivered from said high pressure
centrifugal compressor to said combustion means.
5. The twin spool gas turbine engine as in claim 1 wherein said
first and second flow interconnection means structurally support
and orient said low pressure spool spool to be angularly offset
from said high pressure spool supplying the external load by about
90.degree..
6. The twin spool gas turbine engine as in claim 1 wherein said
combustion means includes a low NO.sub.x annular combustor having
an angled internal cone.
7. The twin spool gas turbine engine as in claim 1 wherein said low
pressure compressor means includes a single entry centrifugal
compressor and said low pressure turbine means includes a mixed
axial/radial flow turbine.
8. The twin spool gas turbine engine as in claim 1 wherein said low
pressure compressor means includes a single entry centrifugal
compressor and said low pressure turbine means includes an axial
turbine.
9. The twin spool gas turbine engine as in claim 1 wherein said low
pressure compressor means includes an axial compressor and said low
pressure turbine means includes an axial turbine.
10. The twin spool gas turbine engine as in claim 1 wherein said
external load driving means is operatively connected to said high
pressure spool shaft means proximate said high pressure centrifugal
compressor.
11. The twin spool gas turbine engine as in claim 1 wherein said
first flow interconnecting means includes a manifold for receiving
said precompressed air from said low pressure compressor means,
said manifold having an increasing cross-sectional area in the flow
direction for diffusing the compressed air.
12. The twin spool gas turbine engine as in claim 1 wherein said
second flow interconnecting means includes diffuser means to
receive and diffuse partially expanded combustion gases exhausted
from said high pressure turbine prior to delivering the gases to
said low pressure turbine means.
13. The twin spool gas turbine engine as in claim 1 further
including bleed means associated with said first flow
interconnecting means.
14. The twin spool gas turbine engine as in claim 1 wherein said
radial in-flow turbine is sized to directly drive an external load
of greater than or equal to about 100 KW and less than or equal to
about 1000 KW.
15. The twin spool gas turbine engine as in claim 1 wherein said
radial in-flow turbine is uncooled.
16. The twin spool gas turbine engine as in claim 1 wherein the
expansion ratio across said radial in-flow turbine is greater than
the expansion ratio across said low pressure turbine means.
17. The twin spool gas turbine engine as in claim 1 wherein the
overall engine pressure ratio is between about 12/1 and 21/1.
18. A high efficiency gas turbine engine for direct-driving an
external load, the engine comprising:
a radial in-flow turbine for operation at a high-peripheral speed,
said radial in-flow turbine having an inlet and an outlet;
means for supplying high temperature, high pressure gases
corresponding to an overall engine pressure ratio greater than
about 12/1 to said radial in-flow turbine, said supply means
including a low pressure rotating air compressor, a high pressure
centrifugal air compressor flow connected to said low pressure
rotating air compressor, and means for combusting fuel with
compressed air from said high pressure centrifugal air compressor
and delivering the combustion gases and all the combustion heat
energy to said radial in-flow turbine inlet under all engine load
conditions;
means for loading said radial in-flow turbine including first shaft
means for directly driving both said high pressure centrifugal air
compressor and essentially the entire external load by said radial
in-flow turbine; and
low pressure turbine means for receiving gases from said radial
in-flow turbine outlet, for extracting work therefrom, and for
driving said low pressure rotating air compressor, said low
pressure turbine means including a low pressure turbine and second
shaft means directly connecting said low pressure turbine and said
low pressure rotating air compressor, wherein said first shaft
means and said second shaft means are rotationally independent for
providing optimum aerodynamic efficiencies for the respective
turbines and compressors associated therewith, thereby providing
high engine efficiency, and wherein the expansion ratio across the
radial in-flow turbine is greater than the expansion ratio across
the low pressure turbine.
Description
BACKGROUND OF THE INVENTION
This invention relates to gas turbine engines using radial turbines
and centrifugal compressors in high pressure engine portion and
arranged so as to achieve low fuel consumption.
Up till now the reduction in fuel consumption in smaller gas
turbine engines has been limited by the difficulty in obtaining
high pressure ratio, cooling small turbine components, and avoiding
parasitic losses resulting from scaled down components, and by the
high cost of miniature precision parts. Had it not been for the
reasons given, small turbine engines would have seen an extended
use generally, particularly because of the ability of such engines
to operate with substantially lower emissions than gas engines,
diesel engines and gasoline engines.
Many previous attempts have been made to solve the above problems,
but the only other solution which addresses the situation uses a
double entry first stage compressor in a single spool gas generator
configuration. This configuration holds promise in the higher power
range, but would be far too expensive for use in the lower power
range, i.e. .ltoreq. about 1000 KW.
SUMMARY OF THE INVENTION
As a consequence of the above mentioned shortcomings in
conventional and known gas turbine engine designs, it is an object
of the present invention to provide highly efficient gas turbine
engines for use in the lower power ranges where fuel consumption
levels and first cost so far have not made them acceptable.
It is a further object of the present invention to provide a gas
turbine engine construction having the flexibility to be utilized
for varying operating conditions without penalizing design point
performance.
In accordance with the present invention, as broadly disclosed and
claimed herein, the high-efficiency, twin spool, gas turbine engine
comprises a high pressure spool having shaft means, and
specifically a centrifugal compressor and a radial in-flow turbine
mounted on the shaft means, the radial in-flow turbine having an
expansion ratio of about 4/1 to 8/1. Means are provided connected
to the shaft means for driving an external load, and combustion
means associated with the high pressure spool are provided for
receiving compressed air from the centrifugal compressor,
combusting fuel using the compressed air, and delivering the
combustion gases to the radial in-flow turbine. The twin spool gas
turbine also comprises a low pressure spool having shaft means and
associated low pressure compressor means and low pressure turbine
means, the low pressure spool being rotatably independent from the
high pressure spool and being essentially load-free. Still further
the twin gas spool-turbine comprises first means for flow
interconnecting the low pressure compressor means and the high
pressure centrifugal compressor for supplying precompressed air
thereto, and second means for flow interconnecting the high
pressure radial in-flow turbine and the low pressure turbine means
for supplying partially expanded combustion gases thereto.
Preferably the pressure ratio of the low pressure compressor is
less than or equal to the pressure ratio of the high pressure
centrifugal compressor, and the low pressure compressor is
preferably, but not necessarily, a single entry centrifugal
compressor and the low pressure turbine can be an axial, radial or
a mixed axial/radial flow turbine. The design load deliverable by
the external load driving means is preferably .ltoreq. about 1000
KW.
The twin spool gas turbine engine may include intercooler means
associated with the first flow interconnecting means for cooling
said pre-compressed air. Preferably the first flow interconnecting
means includes diffuser means to receive and diffuse partially
diffused compressed air received from the low pressure compressor
means prior to delivering the air to the high pressure centrifugal
compressor. Also, the first and second flow interconnecting means
may structurally support and orient the first and second spools to
be angularly offset from one another.
The twin spool gas turbine engine may further include recuperator
means for extracting heat values from the combustion gases
exhausted from the low pressure turbine and preheating the
compressed air delivered from the high pressure centrifugal
compressor prior to combustion, and the combustion means may
include a low NO.sub.X combustor of the angled internal cone
type.
The accompanying drawings, which are incorporated in and constitute
a part of this specification, illustrate a preferred embodiment of
the invention and, together with the description, serve to explain
the principles of the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
A better understanding of the present invention will be had upon
reference to the following detailed description when read in
conjunction with the accompanying drawings in which:
FIG. 1 is a plan cross-sectional view of one embodiment of the high
efficiency twin spool, gas turbine engine according to the present
invention;
FIG. 2 is a schematic view of the gas turbine engine depicted in
FIG. 1 but with intercooler and recuperator components shown;
FIGS. 3A and 3B depict alternative in-line and parallel spool
configurations, respectively, for the gas turbine engine shown in
FIG. 1;
FIG. 4 is a cycle schematic showing the results of a calculational
Example of a twin spool gas turbine engine in accordance with the
present invention; and
FIG. 5 is a plan cross-sectional view of variations of the gas
turbine engine depicted in FIG. 1.
Reference will now be made to the present preferred embodiment of
the invention which is illustrated in the accompanying drawing.
DESCRIPTION OF THE PREFERRED EMBODIMENT
With initial reference of FIG. 1 there is shown a preferred high
efficiency, twin spool, gas turbine engine constructed in
accordance with the present invention and designated generally by
the numeral 10. Modifications and variations to the presently
preferred embodiment depicted in FIG. 1 would be immediately
evident to one skilled in the art after reading the detailed
discussion below in conjunction with the accompaning drawing, and
these modifications and variations are considered part of the
present invention which is intended to be limited only by the
appended claims and their equivalents.
In accordance with the present invention, the gas turbine engine
includes a high pressure ("H.P.") spool having shaft means and
characterized by a centrifugal compressor and especially a radial
in-flow ("R.I.F.") turbine mounted on the shaft means. As embodied
herein and with continued reference to FIG. 1, gas turbine 10
includes H.P spool 11 including shaft assembly 12, and with single
entry centrifugal compressor 14 and R.I.F. turbine 16 mounted at
opposite ends thereof. R.I.F. turbine 16 is configured to operate
with an expansion ration of from about 4/1 to 8/1, and generally
will have a rotor diameter of .ltoreq.500 mm at load levels less
than 1000 KW.
Further, in accordance with the invention, means connected to the
shaft means are provided for driving an external load. As embodied
herein, external load driving means is shown schematically at 18
connecting load 20 to H.P. spool shaft assembly 12 proximate H.P.
compressor 14. Load 20 (which is also shown schematically) can be
for instance a high speed electrical generator. Load driving means
18 can include gear reduction apparatus, clutches, etc. with
related shafting. Driving means 18 can include means for driving
the H.P. spool lubrication system, fuel delivery systems (both not
shown) and other auxiliary systems of engine 10, although these
systems are not considered part of the external load, or the
auxiliary loads can be carried by the low pressure ("L.P.") spool
to be discussed hereinafter, or shared between the H.P. spool and
the L.P. spool.
Further in accordance with the present invention, combustion means
are provided for receiving compressed air from the H.P. centrifugal
compressor, combusting fuel using the compressed air, and
delivering the combustion gases to the high pressure R.I.F.
turbine. As embodied herein, and with continued reference to FIG.
1, gas turbine 10 includes combustion means 22 including combustor
23 which receives compressed air from H.P. centrifugal compressor
diffuser 24 and fuel through conduit 26 from a fuel supply (not
shown). Combustion gases exit combustor 23 and are delivered to the
R.I.F. turbine inlet 28 for subsequent expansion. It is highly
preferred that combustor 23 utilize the low NO.sub.X angled
internal cone annular combustor configuration disclosed in the
present inventor's copending application Ser. No. 07/488,136 filed
concurrently herewith and entiled LOW EMISSIONS GAS TURBINE
COMBUSTOR the disclosure of which is hereby incorporated by
reference.
Still further in accordance with the present invention, the gas
turbine includes a low pressure ("L.P.") spool having shaft means
and associated L.P. compressor means and L.P. turbine means. As
embodied herein, gas turbine 10 includes L.P. spool 30 having L.P.
shaft assembly 32, with L.P. compressor 34 and L.P. turbine 36
mounted on opposite ends thereof. In the preferred embodiment L.P.
compressor is shown as a centrifugal compressor and L.P. turbine as
a mixed axial/radial turbine. Other configurations are possible,
including using pure axial and pure radial compressor and turbine
components.
Importantly L.P. spool shaft assembly 32 is rotatably independent
from H.P. spool shaft assembly 12 such that it can operate at a
variable speed, depending upon the power level, and hence mass flow
requirements, of the operating conditions while allowing the H.P.
spool 11 to continue to operate at essentially constant speed which
is advantageous for constant frequency load demands. The H.P spool
could also be used in a variable RPM application. At low power
conditions, the L.P. spool would slow down lowering the energy
consumption and thereby reduce efficiency-fall off.
Also, L.P. spool 30 is free from external loads. That is, the power
extracted by L.P. turbine 36 is used solely to power L.P.
compressor 34, and possibly to supply power to drive some or all of
the engine auxiliary systems. As discussed previously, these
auxiliary systems are not considered part of the external load
which is driven solely by H.P. spool 11.
Also, L.P. spool preferably includes diffuser 38 for receiving and
finally diffusing the combustion gases exhausted from L.P. turbine
36. As shown schematically in FIG. 1 and FIG. 2, gas turbine 10 can
optionally include recuperator means 40 which can extract heat
values from the diffused exhaust from L.P. turbine 36 and heat the
compressed air from H.P. compressor 14 prior to admission to
combustion means 22.
Further in adccordance with the invention, first and second flow
interconnection means are provided. The first flow interconnection
means channels the pre-compressed air from the L.P. compressor
means to the H.P. centrifugal compressor, and the second flow
interconnection means channels the partially expanded combustion
gases from the H.P. radial in-flow turbine exit to the L.P. turbine
inlet.
As embodied herein, first flow interconnection means 41 includes
manifold 42 for collecting the pre-compressed air from L.P.
compressor diffuser 44 and conduit 46 for carrying the collected
pre-compressed air from manifold 42 to H.P. centrifugal compressor
inlet plenum 48. Preferably, low pressure compressor diffuser 44 is
configured to only partially diffuse the compressed high velocity
air exiting low pressure compressor 34 to a higher pressure and
lower velocity, and that collection manifold 42 is configured with
an increasing cross-sectional flow area in the flow directions
(shown by arrows) to accomplish further diffusion. Because the
outer dimensions of the L.P. module can be influenced by the outer
diameters of the L.P. compressor diffuser which depends, in turn,
on the flow rate and average velocity, accomplishing part of the
diffusion in manifold 42 of interconnecting means 41 can
advantageously achieve a reduction in the engine space
envelope.
Optionally, and as seen in FIGS. 1 and 2 FIG. 2, first flow
interconnection means 41 can include intercooler 50 for increasing
the density of the pre-compressed air prior to final compression by
H.P. centrifugal compressor 14. For operation with intercooler 50,
the pressure ratio of L.P. compressor 34 should be less than or
equal to the pressure ratio of H.P. centrifugal compressor 14. Flow
interconnection means 41 can also include bleed valve 52, shown
connected to conduit 46 in FIG. 1, for use during start-up or other
transient conditions.
As embodied herein, second flow interconnection means 54 includes
diffuser 56 for diffusing the partially expanded combustion gases
exiting H.P. turbine 16 and manifold 58 for distributing the
diffused, partially expanded gases to the L.P. turbine inlet 60. In
the preferred gas turbine engine 10, the L.P. spool pressure vessel
is configured to provide manifold 58. Diffuser 56 preferably can
include vanes 62 to eliminate swirl from H.P. radial turbine 16
exhaust.
First and second flow interconnection means 41 and 54 also
structurally support and orient H.P. spool 12 to be angularly
offset from L.P. spool 30. The offset shown in FIG. 1 is about
90.degree. but other angles may be preferred depending upon the
intended application, and even the zero offset angle configurations
shown in FIGS. 3A and 3B. The 90.degree. configuration shown in
FIG. 1 minimizes heat and flow (form and friction) losses between
the H.P. spool and L.P. spool and L.P. turbine to the atmosphere
and is thus preferred from increased thermal efficiency
considerations and mechanical convenience.
FIG. 5 shows several variations in the configuration of low
pressure spool from that shown in FIG. 1. In FIG. 5, the low
pressure spool (designated 30') of engine 10' includes low pressure
turbine 36' which is configured as an axial turbine, compared to
the mixed axial/radial turbine 36 shown in FIG. 1. Axial low
pressure turbine 36' drives low pressure centrifugal compressor 34
via shaft assembly 32 in the FIG. 5 depiction. However, an axial
compressor component such as axial compressor 34' (shown
schematically in dotted lines) could be substituted for centrifugal
compressor 34 in L.P. spool 30', in accordance with the present
invention. Importantly, for all variations shown in FIG. 5 and
described herein, H.P. spool 11 remains configured with a
centrifugal compressor such as 14 and a radial in-flow turbine such
as 16.
As is evident from the above description, the present invention
relates to a gas turbine engine with two separate shafts each with
associated compressor and dedicated turbine components. Contrary to
conventional configurations, the power output is taken essentially
exclusively from the high pressure part of the engine. This
invention relates only to turbine engines where the high pressure
turbine is of the radial type, where the high pressure compressor
driven by the said radial turbine is a centrifugal compressor, and
where the external load is driven by the H.P. spool. See FIG.
1.
The two shafts with their corresponding rotor components are called
"spools." The high pressure spool therefore has a centrifugal
compressor and a so-called radial in-flow turbine.
In the preferred embodiment, power is extracted from the compressor
side of the H.P. spool, but a power take-off from the turbine side
should also be considered within the scope of the invention as long
as power is extracted from the H.P. spool.
The sole purpose of the L.P. spool is to absorb additional energy
from the exhaust gas after it has expanded through the H.P. turbine
and to use this energy to drive the compressor which is mounted on
the same L.P. shaft. This compressor, in turn, provides compressed
air (or other gaseous medium able to support combustion) to the
second stage compressor so as to reach the designated overall
pressure ratio for the engine through the use of both compressors,
typically, about 20/1.
No external power will be extracted from the L.P. spool. Auxiliary
drives, such as fuel and oil pumps, etc. are not considered as
external power in the context of this invention and could be driven
by the L.P. spool, the H.P. spool, or shared.
The purpose of the invention with the incorporation of a radial
in-flow turbine as the driving component for both the H.P.
compressor and the external load is to provide the maximum possible
work for the radial in-flow turbine without exceeding the
mechanical properties in the turbine caused by the resulting high
peripheral speed of the turbine rotor. This higher peripheral speed
is made possible because of the "Eiffel tower" profile of the
radial turbine blade with a thick section in the hub-region and a
thin, lighter section at the blade tips. An axial turbine does not
have this feature and would, in contrast to a radial stage, also
incur a large loss in efficiency at high expansion ratios. Also the
resulting high peripheral speeds of the radial turbine lowers the
blade temperature thus making blade cooling unnecessary at
temperature levels where axials need it, an important feature for
small turbine engines.
Reverting to FIG. 1, the flow path of the engine is described
below: Air enters the low pressure compressor 34, is compressed and
exits via manifold 42 into the transfer conduit 46 leading to the
high pressure compressor 14. After having been further compressed
by centrifugal compressor 14 the air enters the combustor 23 where
it is burned with fuel added. The resulting hot gases are expanded
through high pressure radial in-flow turbine 16, which in turn
drives both the high pressure compressor 14 and the external load
20 via the shaft 12 and driving means 18. After the gases have
passed the H.P. turbine 16, further exhaust energy is recovered in
an exhaust diffuser 56, here shown as a radial type to remove
"swirl" from the exhaust with minimum losses. In the preferred
embodiment the L.P. shaft 32 and the H.P. shaft 12 are located at
an angle with respect to each other (here shown as 90 degrees). The
gases are only partly expanded in the H.P. turbine and enter the
L.P. turbine via manifold 58. Following final expansion through
L.P. turbine 36, shown in FIG. 1 as a "mixed flow" type, the gases
are exhausted through diffuser 38.
The incorporation of intercooler and recuperator components in the
flowpath such as intercooler 50 and recuperator 40 can be
conveniently effected with a very minimum of changes, and thus
achieve the arrangement described in FIG. 1 and FIG. 2. However the
intercooler and recuperator components are not essential parts of
the present invention.
Since the difference between the turbine inlet temperature (i.e.
approximately the average combustor exit temperature) and the
relative stagnation temperature for a typical radial turbine stage
at the inlet will be nearly 50% of the total temperature drop over
that expansion stage, a higher expansion ratio results in a lower
turbine blade temperature. This feature allows higher cycle
(combustor) temperatures to be utilized for increased thermal
efficiencies.
The best state of the art radial turbine will provide good
efficiencies up to about 8/1 expansion ratio. This would correspond
to approximately the same pressure ratio for a single shaft radial
engine. A simple cycle engine, i.e. an engine without recuperator
or other cycle additions to increase thermal efficiency, will need
substantially higher pressure ratios in order to achieve the best
efficiencies for a given turbine inlet temperature of say,
1000.degree. C. The current invention provides increased efficiency
at the design point by utilizing a higher overall cycle pressure
and, hence, improving shaft efficiency by means of the
pre-compression generated by the L.P. spool.
According to the invention, the largest part of the expansion
should be taken over the H.P. radial turbine in order to obtain the
above mentioned effect of minimizing turbine blade temperatures for
a given average combustion temperature. For about a 21/1 pressure
ratio engine the expansion ratio over the H.P. turbine would be
typically about 7/1 and over the L.P. turbine about 3/1. However,
overall pressure ratios of between about 8/1 and 30/1 are
contemplated, with the higher pressure ratios corresponding higher
outputs. For power output in the range below 1000 KW, the over all
pressure ratio will generally range between about 12/1 and
21/1.
Component efficiencies of centrifugal compressors and radial
turbines are very dependent on Specific Speed. The term Specific
Speed (N.sub.s) defines the speed-flow-work relationship in a
component and is described in U.S. Pat. No. 4,641,495 in
conjunction with another engine configuration of the same
inventor.
In conventional gas turbine engines with centrifugal and radial
turbines, an acceptable N.sub.s is obtained either at the expense
of pressure ratio, or by using a double entry first stage
compressor in the gas generator, a configuration disclosed in U.S.
Pat. No. 4,461,495. In such configurations, a separate power
turbine must be used to take out external shaft power. This last
arrangement is significantly more complicated than the current
invention which is therefore more suitable for the lower power
range.
The engine configuration described herein has two rotatably
independent shafts which operate at different speeds so as to
satisfy the requirement for an optimum specific speed for the high
pressure spool wheels and at the same time provide the best
possible flow conditions for the L.P. compressor and turbine.
The engine configuration according to the invention requires that
the power output is taken from the high pressure spool which is
also generally the highest rpm shaft. This feature could prove to
be a minor disadvantage if a reduction gear is contemplated as part
of load driving means 18, as a somewhat greater reduction ratio
would be required. Compared with the significant advantage of being
able to operate at a H.P. turbine inlet temperature of some
100.degree. C. higher than an uncooled axial turbine can accept,
this should not detract substantially from the value of the
invention. Current developments in the field of direct driven high
speed electric generators would fit ideally into an energy system
with the above-described engine.
The torque characteristics of the engine configuration will be
unique in that it will lie between a single shaft engine
characteristic and the characteristic of an engine with a free
power turbine. The reason for this is the nature of the
above-described engine as a "boosted" single shaft machine.
Depending on the application a certain degree of variability in air
flow would be required between and/or possibly within the two
compressor and turbine stages. The simplest construction for
accommodating for such variability takes the form of an interstage
bleed, such as bleed means 52 shown in FIG. 1, to be activated
during transients to help match the flow to the components.
Both a relatively high pressure cycle and a lower pressure ratio
engine can be constructed according to this invention. The lower
pressure ratio version would normally be equipped with a
recuperator and an intercooler between the compressor stages. The
two separate compressors and the construction of flow
interconnecting means 41 make the adaptation of such components
very convenient, see FIGS. 1 and 2.
The mechanical arrangement of the engine can take several forms in
addition to the angularly offset spool configuration shown in FIG.
1. FIG. 3a shows an in-line spool configuration, and FIG. 3b
depicts schematically a parallel spool arrangement. These latter
configurations are considered to fall within the general scope of
the invention.
With similar component technology and turbine life, an engine
according to the invention of some 400 KW which would achieve 30%
thermal efficiency would compare favorably with 22% for a
conventional design--both configured as "simple cycle" engines. For
recuperated engines, the comparison would be about 40% and 35%,
respectively.
A typical conventional design would have a single spool with a
single stage compressor in this power range. No provisions for an
intercooler can then be made. Consequently the required mass flow
and hence the recuperator would be some 30% larger in volume.
The efficiency advantage of the disclosed design will be the
greatest between about 100 KW and 1000 KW. Larger than 1000 KW
could be considered, however, but the manufacturing cost of large
radial wheels and their corresponding stators would eventually
limit its use. Conversely, in the small sizes, the manufacture of
low cost radial and centrifugal components is very well established
in the production of turbochargers. The component efficiencies in
these kinds of wheels are also intrinsically higher than for small
axial-type compressor and turbine stages.
Twin spool engines normally have a separate power turbine on a
separate shaft. The Rolls-Royce Tyne turboshaft/turboprop is an
exception to this with the power take-off from the low pressure
spool. Unlike such previous engine designs, the configuration
according to the present invention extracts power from the high
pressure spool with the express purpose of loading the radial
in-flow turbine as much as possible and thereby utilize its
characteristics of lowering the metal temperature of the wheel for
a given combustion average temperature. Combining this feature with
the previously mentioned advantage of a twin spool operation
provides both additional temperature capability and higher pressure
ratios, key factors to achieving good cycle efficiency.
The H.P. spool shall, according to the invention, always use a
centrifugal compressor and a radial in-flow turbine. The L.P. spool
could have components of both axial, mixed flow or
radial/centrifugal type, with the latter being more suitable for
the lower end of the contemplated power range.
The gas turbine of the present invention can have several
configurations, among these, the following are mentioned:
(A) Simple cycle engine without intercooler or recuperator.
(B) Simple cycle engine with intercooler to improve power output
and some efficiency improvement as well. This is at some expense of
added complexity. Applications desiring combined power and heat
(steam) would advantageously use configurations (A) or (B).
(C) Same as (B) but with the addition of a recuperator. This gives
great flexibility in the market and will, for the recuperated
version, provide substantial efficiency improvement. The added cost
and complexity with the recuperator will be substantial,
however.
(D) Similar to (C), but with component/cycle matching made optimal
for the recuperated version and the achievement of the highest
possible efficiency. When this version is operated without
recuperator, but with intercooler, the efficiency will not be quite
as good as (B). Base load electric power w/o any requirement for
heating favours this optimized recuperated cycle.
EXAMPLE 1
In order to better illustrate the present invention, a
calculational example is provided for a nominal 400 KW application
using a twin spool gas turbine engine with intercooler.
FIG. 4 presents the results of the calculations and shows that an
overall efficiency of over 30% can be expected for the depicted
simple cycle with intercooler ("SCI"), based on a combustor exit
temperature of 1051.degree. C.
The cycle shown in FIG. 4 and described below uses the same engine
hardware for an engine both without and with a recuperator and
achieves, as a consequence, a very desirable market
flexibility.
If this cycle is turned into a Recuperated Cycle (RC), simply by
adding a recuperator to the same engine hardware and making minor
adjustments, the performance will be as follows:
______________________________________ Power output 375 KW Shaft
Thermal Efficiency 40.1% ______________________________________
The component assumptions for the SCI and RC are as follows:
______________________________________ Overall Pressure Ratio 15/1
L.P. Comp. EFF. Isentropic 82.8% H.P. Comp. EFF. Isentropic 80.4%
Overall Expansion Efficiency Total-Stat. Isen. 87.0% Recuperator
Effectiveness 85% ______________________________________
EXAMPLE NO. 2
For comparison purposes, the calculations on the FIG. 4 cycle were
repeated for the case of a lower combustor exit temperature. An
initial rating using a 50.degree. C. lower combustor exit
temperature than the 1051.degree. C. target rating level would
yield 30.1% efficiency and 363 KW for the SCI, if the engine was
matched for this situation. Under the same circumstances the RC
would give 38.3% efficiency and 341 KW.
It will be apparent to those skilled in the art that various
modifications and variations can be made in the above-described
embodiments of the present invention without departing from the
scope or spirit of the invention. Thus, it is intended that the
present invention cover such modifications and variations provided
they come within the scope of the appended claims and their
equivalents.
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